Reactive propulsion power plant having radial flow compressor and turbine means



May 31, 1949. N c PRICE REAGTIVE PROPULSIdN PbWER PLANT HAVING RADIAL FLOW COMPRESSOR AND TURBINE MEANS Filed Feb. 14, 1944 4 Sheets-Sheet 1 IN VEN TOR. MTHAN GPRIDE AGENT 1949- N. c. PRICE 2,471,892

REACTIVE PROPULSION POWER PLANT HAVING RADIAL FLOW COMPRESSOR AND TURBINE MEANS Filed Feb. 14, 1944 4 Sheets-Sheet 2 LEFT HAND THREADfi. 43 30 2a 45 49 as I INVENIOR. MTA 'AN CPRIL'E May 31, 194 2,471,892

N. 0. PRICE REACTIVE PROPULSION POWER PLANT HAVING RADIAL FLOW COMPRESSOR AND TURBINE MEANS Filed Feb. 14, 1944 4 Sheets-Sheet 3 GENERATOR FUEL PUMP 14a INJECTION AIR BOOSTER PUMP "a 0 [A35 H9 I 1 INVENTOR.

.Mu'w 5. PRICE N. C. PRICE REACTIVE PROPULSION POWER PLANT HAVING RADIAL May 31, 1949. 2,471,892

FLOW COMPRESSOR AND TURBINE MEANS 4 Sheets-Sheet 4 Filed Feb. 14, 1944 INVENTOR. NATHAN C. PRICE Agent A Patented May '31, 1 949 'REACTIV E PROPULSION POWER PLANT HAV- ING RADIAL FLOW COMPRESSOR AND TURBINE MEANS Nathan 0. Price, Hollywood, Calif assignor to Lockheed Aircraft Corporation, Burbank, Calif.

Application February 14, 1944, Serial No. 522,342

This invention relates in general to prime movers of the fluid reaction type and more particularly to the type of internal combustion engines which function in accordance with the principle and in the manner commonly known asfjet propulsion. o o o The features of this invention may have numerous applications but they find their principal application in prime movers for aircraft and the like high speed vehicles, particularly for the propulsion of airplanes at high velocity.

The constant pressure fluid reaction type of internal combustion units heretofore known have usually been constructed along the general lines of the larger stationary types of prime movers used for generating power. The gas turbines and compressors of such units have been widely separated from one another and interconnected by means of relatively long shafts and more or less complicated transmission gearings with their attendant relatively low mechanical efficiency and heat dissipation problems. The prior. art units usually employ axial flow turbines or when employing radial flow turbines the movable blade rows have all been carried on a common rotor and thus allrotated at the same angular speed. Where all of the blades are carried on a common rotor,

conditions of maximum eiiiciency can not be maintained at all radial regions within the expansion or compression zone, as the case may be, because ofthe difference in velocity of the blade elements at such different zones. The best construction under such conditions must necessarily be a compromise. which is not so eflicientas could .be desired. Additionally, in such conventional vide a gas turbine and compressor combination of improved and novel design wherein the thermal and mechanical efliciency is substantially improved by the elimination of substantially all of the transmission shafting and gearing of the 21 Claims. (01. .6035.6)

type heretofore employed in connection with, turbo compressor units of this type employed for the production of a reactive fluid jet. o s It is-a still further object of this invention to provide a simplified form of turbo-compressor rotor and blading which lends itself to simple and inexpensive production methods.

s The objects of the invention are also attained by a compact supercharger compressor rotor arrangement wherein a plurality of concentrically positioned radially interconnecting squirrel-cage ty e rotors carrying aplurality of peripherally spaced, longitudinally positioned impeller blades, are arranged for radial flow of the propulsive gases therethrough, eachof such squirrel-cage rotors being freely and independently rotatable counter to one another.

It is a still further object of this invention to provide a turbo-compressor in which the. blade velocities of each stage may vary independent of one another whereby the most efficient blade angles of incidence may be employed throughout the several stages, the compression ratio maintained substantially linear throughout the stages and blades of equal length throughout the several stages may be employed, all of which contributes to the improved efficiency of operation and simplification of construction of the machine. It is likewise apparent that the suitability of blades of equal length, and the suitability of squirrel-cage rotor blades which do not require twist from one end,of the blade with respect to the other to attain correct angles ofincidence, greatly facilitates ease of manufacture and reduces cost. Furthermore, the elimination of requirement of blade twist relieves the difllculty of initial development of suitable blade forms, since the prediction of pressure distribution and angle of incidence spanwise along the blade does not vary. Unlike conventional constructions, entailing twisted blades, the same airfoil section may be employed along the entire blade length.

The objects of this invention are attained, in general, by providing a power plant employing a turbo-compressor unit of high efliciency and extreme compactness in which substantially all shafting and gearing have been eliminated, and the structure in general simplified and which produces propulsive force andwork wholly by means of the reaction of a high velocity expandable fluid jet. The attainment. of the objects and: advantages of this invention are realized largely by an improved design of the turbosupercharger combination wherein each of-the pluralities. of supercharger compressor cascades is individually and directlycoupled to anddriven from each of the plurality of turbine cascades cades directly driven and independently rotatable results in improved thermal and mechanical efficiency and weight reducing and power saving elimination of inter-coupling drive mechanisms and other associated mechanical equipment heretofore thought to be indispensable.

walls 14' and I which are coupled together at the outer housing joint X. The forward annular pas- These and other objects and features of novelty will become evident hereinafter in the description which, together with the following drawings il-v 4-4 of Figure 3.

Figure 5 is a typical fragmentary cross escotional view of the turbine blades taken on line 5 -5 of Figure 3.

Figure 6 is a front end elevation of the unit with the accessory compartment cover removed.

Figure 7 is a partial enlarged cross-sectional view of the compressor discharge and turbine inlet nozzle assembly shown in Figure 1.

Figure 8 is an enlarged fragmentary crosssectional view taken on line 0-2 of- Figure 7.

Figure 9 is an enlarged fragmentary detail of a turbo-compressor rotor bearing.

Figure 10 is an enlarged fragmentary longitudinal section of a portion of the combustion chamber illustrating the fuel jet and nozzle means.

Figure 11 is a fragmentary transverse sectional view taken as indicated by line ll-ll on Figure 10. V

Figure 12 is a fragmentary stretch-out view of the swirl vanes shown in Figures 10 and 11. Referring now to the drawings, in which like reference numerals refer to corresponding parts throughout the several figures, the apparatus of the invention is as follows:

The power plant assembly as shown in the drawings comprises nine main components, namely: an outer streamline housing having a forward portion H1 and a rearward portion Hz joined at-X; an inner streamline housing having a frontal supporting structure portion Hz, and a rearward enclosing portion H4, supported coaxially together substantially concentrically within the outer housing, and within the inner housing, a turbo-compressor T having a radial flow compressor section C, and a radial flow gas turbine section G, a centrally located combustion chamber Z interconnecting the compressor and turbine sections, and a forwardly located accessory compartment Aenclosed by the conical cover B.

The inner streamlined housing Ka -H4 which contains the principal components of the apparatus is concentrically supported within the outer housing H1 as before stated, at the forward end by means of a plurality of interconnecting streamlined struts l0 extending between said outer housing H1 and said inner housing portion H: and at the rearward end by means of a plurality of interconnecting streamline struts III, forming therebetween forwardly located. and rearwardly located annular shaped passages l2 and I3 respectively, separated endwise from one another by curved interveningcircumferential sage 12 leads fromthe forwardly directed. inlet Ii rearwardly totthe space H which forms a peripheral inlet of the compressor 0. The rearward annular passageway similarly formed betvflaen the rearward portion of the outer hous- 'ing H2 and the rearward portion of,the inner housing H4, leads from the peripheral-outlet or exhaust receiving space I! of the gas turbine G rearwardly and inwardly to the axial propulsive discharge nozzle N. The housing H1 may be fabricated from chromium steel sheet, while the housings Hz and H; may be formed from chromium-nickel-iron sheet which is heat resistant.

Within theforwardinner housing H: and integral therewith a primary structure is provided for rotationally supporting the elements of the compressor and turbine, for carrying the various accessories and tying together the several housing components of the unit assembly. The structure of the compressor and turbine will be more fully described hereinafter. The inner primary supporting structure comprises a longitudinal disc or diaphragm like member 20 ofsuch material as chromium-nickel-manganese steel heat treated to a tensile strength of 160,000 pounds per square inch, having a plurality of radially extending integral stiffening. webs 2| formed as inward radial extensions of the supporting vanes Ill. The diaphragm 20 together with the stiffening webs 2| which join with or may be formed integrally with the curved surface of the housing H3 at 22 is thus supported by the plurality of radial interconnecting struts III which struts are in turn suitably attached at their outer extremities at 23 to the inner wall of the outer housing structure H1. The inner diaphragm 20 is thus supported as a solid unit by means of the supporting vanes ll, webs 2|, and the inner and outer surfaces 22 and 23 of the housings H: and H1. 7

The rearward face of the vdiaphragm 2| carries a plurality of circular recesses as best shown at 25 in Figure 3 forming thereby a plurality of concentric ring shape tracks 26 for rotationally supportingthe concentric turbine blade rings or cages which are hereinafter more fully'described.

Each turbine-ring-supporting-track 26 carries in its outwardly facing cylindrical surface a plurality of spherical seats 21 in each of which is seated the spherical supporting pedestal 20 of a pivotable slipper-type bearing 29 as best shown in Figures 3 and 9. Pivotally carried by the spherical supporting pedestal 20, is a segmental bearing element ll having an outer cylindrical bearing surface 3| conforming to and adapted to support the inner cylindrical bearing surface of the turbo-compresso more fully described.

The turbo compressor portion T of the unit, rotatably supported by means of the diaphragm 20 comprises a plurality of concentric, radially interconnecting,-rotors I2 to 21. Each rotor comprises a plurality of circumferentially spaced airfoil shaped compressor impeller blades 39 and airfoil shaped turbine blades 40 axially positioned and coaxially joined end to end'by means of an intermediate ring or separator II. The outer end of the compressor impeller blade 39 and the turbine blades 40 terminate in end rings 42 and 43 respectively. The unitary structure formed by the end rings 42 and I3 and the ring ll joined together by the plurality of compressor and turbine blades 39 and 40 respectively cage rlmner ashereinafter intermediate drical rows of impeller and compressor blades as best shown in Figures 3, 4 and 5. The impeller blades in alternateconcentric rows have opposite angles of incidence as best shown in Figures 2, 4 and 5 adapting alternate rows thereof for counter-rotation with respect to one. another. The

size and positioning of blading of the several concentric rotors are preferably constructed so that their solidities throughout are substantially equal, that is the solidities of the rotors from the innermost to the outermost one is substantially constant. By'the term solidity is meant the ratio of the chord length of the impeller blade to its circumferential arc spacing center to center or leading edge to leading edge approximately.

' The forward faces of each of the rotor rings 42 adjacent the supporting diaphragm 20 is provided with a unitary annular runner 45 carrying on its inner cylindrical surface a high polished annular groove or recess 46 of such a width and depth as to adapt it to receive the segmental bearing elements 30 of the slipper bearings as shown at 29... Both axial and radial forces may thus be taken by the segments 38 of the bearings 29 acting against the bottom and the sides 41 and 48 of the. groove or recess formed in the runner member 45. The. segmental bearing elements 38 are generally composed of a mixture of bronzeand iron powder molded under high hydraulic pressure to have microscopic pores for retaining oil. A very light lubricant of approximately Saybolt seconds viscosity is suitable for these bearings which are designed to operate at high rubbing speed. Annular packings 49 and 58 are provided in suitable grooves of the faces of the bearing supporting track members 26 and are adapted to make a sliding gas tight seal with the adjacent shoulders 5| and 52 of the rotor rings which shoulders are formed on either side of the annular rotor supporting runner 45 as best shown in Figure 3. The packings 49 are generally composed of a mixture of carbon and silver powder molded under high hydraulic pressure to form a heat resistant material with a very low coefficient of friction. The adjacent cylindrical surfaces of the concentric central rings 4| of the turbine rotors are provided with a plurality of circum ferential serrations or threads 53 which are adapted to impede or control longitudinal flow of gases through the clearances between the rings, from the compressor section to the turbine section or vice versa, depending upon the pressure gradient. Threaded serrations are preferably employed. with adjacent threads on opposite rotor rings having pitches of opposite hand and having leads with respect to the direction of rotation such as to tend to move the gases therein counter-current to the leakage pressure gradient. For example, in Figure 3 it is assumed that the leakage pressure gradient is from the turbine toward the compressor in which case the threadlike serrations are pitched as indicated by the legends in that figure.

The compressor impeller blades 39 and the turbine impeller blades 48 are made hollow for lightness. The rotorsand blading may be constructed of an alloy high in nickel, chromium and cobalt with small additions of other metals such as tungsten, columbium, copper and aluminum. Such alloy has very high yield strength and favorable creep characteristics through a wide temperature range. The hollows of each of the turbine impeller blades 48, as shown in Figure 3, are provided with a transverse bafiie or partition .55 which divides the hollow into parallel channels 58 and 51 joined at their inner ends at 58. Com municating with the channels 56 and 51 in the hollow ofthe turbine impeller 40 are a pair of vent'holes 59 and 68 drilled longitudinally through the end rings 43.

Adi acent the innermost turbo-compressor rotor 31 of the plurality of concentric turbo-compressor rotors 32 'to 31 is an inner cage-like nozzle Q assembly comprising the compressor section out-' let nozzle vanes 6|, the turbine section inlet nozzle vanes 62 joined end to end by means of the intermediate hollow separator 83 formed by the bifurcating walls 64 and 65 thereof and terminating at their forward and rearward ends in the forward and rearward chamber walls 68 and 81 respectively. The compressor discharge nozzle vanes BI and the turbine inlet nozzle vanes 62 are hollow as best shown in'Figures '7 and 8 and the hollows of these blades or vanes are interconnected through 83 and communicate with the annular space 68 which is in turn in communication with the annular channel 69. The channel 69 is connected at its outer extremity with a return annular channel 18 which leads to a forwardly directed annular slot H formed through the wall of the combustion chamber discharge passage 12. The compressor discharge nozzle vanes 8| are provided with a plurality of longitudinally spaced perforations I5 as best shown in Figure 7, said perforations being in the leading edge or stagnation point of said vanes and adapted to form intercommunication between the compressor discharge and the compressor discharge vane hollows. Compressed air may thus enter the perforations l5 and flow through the hollows of the vanes BI and 62 into the annular chamber 68 and thence circulate through the passageways 69 and 18 to discharge through the slot ll into the combustion chamber discharge passage 12.

A tube ll passes substantially axially through the hollows of the interconnected compressor discharge vanes and the turbine inlet nozzle vanes through the hollow 63 and serves to interconnect the air passage chamber 68 with the forward annular space 88 ahead of the wall 8| which in turn is in communication with the accessory transmission chamber 82 and the plurality of annular and cylindrical clearance spaces surrounding the bearing supported, forward end of the turbine rotor rings. The tube I! thu serves for axial pressure equalization between the forward and rearward ends of the plurality of concentric turbine rotors. End thrust on the turbine rotors due to external pressure unbalance is thus substantially eliminated.

The inward radial discharge of the compressor section C leads through the beforementioned discharge nozzle vanes 6| through a curved passage 83 to a central, axially extending conduit defined by a coaxial cylindrical wall 84 which passage in turn communicates by way. of bends 85 and 86 containing a plurality of Frey bend vanes 01 staggered turning vanes as shown at distance from the inner surface of the cylindrical walls 84 and 90 respectively to form narrow shroud members 9| and 92 are supported from the inner and outer surfaces 84 and 90 of combustion chamber Z by means of a plurality of circumferentially spaced positioning lugs or struts as shown in Figure 1 at 95 and 98. Within the forward-inlet portionof the combustion zone centrally located between the shroud members 9| and 92 is a relatively short annular member formed with two converging walls 91 and 98 joined in an apex at the forward edge at 99 and opening rearwardly as shown at I to form a fuel .spray nozzle housing. The said walls 91 and 98 of the fuel nozzle housing are supported between and join the inner rearward surfaces of the shrouds 9| and 92 by means of inner and outer concentric sets of a plurality of circumferentially interconnectingswirl vanes I0I and I02. These swirl vanes have an exit angle of about 25 to the tangent and form at the entrance of the combustion zone on either side of the fuel nozzle housing 91-98 between the shrouds 9| and 92, two vaned concentric annular inlet passages which are adapted to impart a spiral fl'ow to fluid flowing to the combustion zone.

A plurality of fuel spray jets extend into the rearward opening portion of the central annular nozzle housing formed between the beforementioned converging walls 91 and 98 and each spray jet carries at the inner end a spray head I03 provided with a pair of laterally directed orifices I04. Each said spray head is supported by and adapted to be supplied with a mixture of liquid fuel-and injection air-by means of a pair of concentric Pipe nipples I05 and I05 which make connection with concentrically arranged circular pipe manifolds I01 and I09 which are contained within the annular entrance space 89 of the combustion chamber Z. The exterior circular manifold I0! is adapted to be supplied with air under suitable pressure through a lead-in pipe I09 and the interior manifold I08 is adapted to be supplied with liquid fuel under pressure through a lead-in pipe H0. Both the air pipe I09 and thehquid fuel pipe Il0 enter the forward end of the combustion zone through the wall III which defines the rearward end of the air conduit 84 and the entrance channel 89 of the combustion chamber Z. The lead-in or supply pipes I09 and III) are supported and passed axially forward through the axis of the conduit 84, through the forward curved walls III. and H3 and coaxially through the bore of the hollow gear hub H4 into the accessory compartment A where it makes connection with a compressed air booster pump as more fully described hereinafter.

The beforementioned converging Walls 9! and 98 of the nozzle housing between which the fuel spray jets extend are provided with apertures II5 registering with each of the beforementioned laterally directed jet head orifices I04 whereby an atomized mixture of fuel and air may be projected laterally into the adjacent space in the contracted portion of the air passages which lead into the combustion zone 93. Y A plurality of jetforations 990 are provided in the apex 99 of the nozzle housing 91-98 for flow therethrough of a relatively small quantity of cooling air.

Under the forward cover B and within the accessory compartment A and supported on the forwardly extending face 3i er. the diaphragm 20 are a number of accessories namely a generator '3, a starter motor I", a liquid fuel pump H8, and an injection air booster pump H9. The

' fuel pump II8, generatorili and injection air pump H9 are gear-driven through a suitable train of gears by the innermost turbo' compressor rotor 31. The shafts of the fuel pump II8. generator H6, and air booster pump II9, pass through the central portion of the diaphragm 20 through suitable bearings at I20 in the manner ,shown in connection with the generator H8 in Figure 1. All of these shafts carry on the inner end thereof, gears as shown at I2I which mesh with a single central pinion I22 which is in turn carried upon the beforementioned central shaft or hub II4 journaled in a central hearing at I23 located at the center of the diaphragm 20. The said hub II4 carries an integral disc member I24 which has formed on its outer periphery a plurality of teeth which are adapted to mesh with a plurality of mating teeth carried on the inner forward surface of the forward turbo-compressor ring at I25 to form a splined connection therewith. The hub H4 and disc I 24 are thus rotatably coupled with the innermost turbo-compres sor rotor 31 yet the splined connection prevents starter shaft I28 carries a toothed clutch member I30. The tooth clutch member I30 is adapted to couple with a corresponding tooth clutch member I3I which is carried on the extending end of the starter motor shaft. The turbo-supercharger is thus adapted to be started by initially rotating the innermost turbo-supercharger rotor by means of the starter motor I I1 acting through the clutch I 30I3I, the gearing I26 and I2! and the spline I25. Upon actuation of the starter motor II! the clutch members I30 and I3I are automatically brought into engagement by conventional means, not shown.

The injection air pipe I09 which enters centrally into the accessory chamber A makes connection with the discharge of the booster air pump II9 through a lateral connection I35. The air inlet to the booster pump II9 makes connection through connecting piping I35 to a forwardly directed inlet conduit I31 which is centrally supported within the open, forwardly directed apex of .theaccessory compartment cover B by means ofa plurality of radial supporting webs as shown at I38. The forwardmost tip of the apex of the accessory compartment cover B is provided with a central opening as shown at I40 adapted to allow rammed air to flow into the accessory compartment in the manner illustrated by arrows Ill and also to flow into the forwardly directed open end I 42 of the beforementioned conduit I31.

Air entering the accessory compartment A through the forward opening I40 is provided with a means of escape through a plurality of channels I45 which pass outwardthrough the supporting webs 2| and I0 and discharge through Q suitable apertures in the trailing edges thereof into the annular channel I1 leading to the peripheral inlet to the compressor section 0. The air entering at I40 and circulating through the accessory compartment A and issuing therefrom through the channels I45 serves to withdraw heat from the turbo-compressor bearings support 20 from the required bearing lubricant used therein and from the accessories, and prevents the accumulation of fumes.

The liquid fuel pipe I I contained concentrically within the air pipe I09 passes therewith into the accessory compartment A and makes connection with the discharge of the fuel pump H8 through a lateral connection I41 passing through the wall of the said air pipe'l09. Liquid fuel is supplied under suitable pressure to the fuel pump H8 by means of a pipe Mil-I49 which enters the accessory compartment through one of the struts III-2| and makes connection with the inlet I50 of the-fuel pump I I8.

The operation is as follows:

. It the unit'is stationary and thus no rammed air is available for entrance through the forward inlet IE to flow through the annular channel I2 into the peripheral inlet of compressor C, the starter motor H1 is actuated to bring the clutch I30-'-I3I into engagement and to drive the innermost turbine rotor 31 to about 10% of full operating speed. The resultant inflow of air radially through the impeller blades of the compressor section C will induce a corresponding rotation of all of the concentric freely rotating turbo- 8 charged from the combustion chamber Z through the annular channel 12 and the inlet nozzle ring.

62 and flow outward radially through the turbine impeller vanes of the plurality of turbo-compress sor rings 32-31. The partially spent gases exhausted from the turbine at I9 are discharged rearwardly through the annular passage I3 formed-between the outer and inner housings Hz The ignition of and H4 respectively and are finally discharged rearwardly in the form of a comparatively-high lowtemperature by virtue of the high inherent velocity propulsive jet expanded to a relatively,

efliciency of the turbo-compressor elements in" When the unit has a relatively high forward airspeed, rammed air enters the forward opening I8 and thereby undergoes a measure of initial compression prior to entrance into the compressor C.

Referring again to the combustion chamber Z and the associated burner construction, the compressed air entering the combustion chamber through the annular entrance space 09, is divided several ways, a. major portion thereof flowing through the two concentric annular inlet passageways containing the swirl vanes IM and I02 formed between the shrouds 9| and 92 and located on either side of the fuel spray nozzle housing 91-98 and thence through the Venturi shaped passageway into the combustion zone 93. These swirl vanes may have an angular position with respect to the axial inlet flow of the air of approximately 25. Another portion of the air from the entrance 89 passes between the combustion zone shrouds SI and 92 and the inner surface of the combustion zone walls 84 and 90 for the purpose of cooling them. Still another minor portion of the air passes around the manifoldpiping" I01I08 and through the passage way formed within the spray nozzle housing 91-98 and thence through the end perforations at 99 into the combustion zone 93.

Under normal operating conditions liquid fuel such as, for example, gasoline, kerosene, or the like, introduced into pipe I I0under pressure from the fuel pump H8, issues from the orifices I04 together with atomizing air introduced into pipe I09 and into the manifold I06 under pressure and passes laterally out through the registering apertures I04 in the-form of a spray into the beforementioned Venturi shaped air passages where it meets andmixes with the flowing air stream, and the resulting fuel-air mixture passes forwardinto the combustion zone 93 where combustion takes place, As before stated combustion may be initiated by a spark plug or glow plug in the manner described in my copending application Serial No. 488,029. The beforementioned swirl vanes IM and I02 at the entrance to the Venturi shaped air passages leading to the com bu'stion zone 93 impart a spiral flow pattern to the burning atomized fuel air mixture flowing through the annular space of the combustion zone. A more thorough mixing of the fuel and combustion air is thus effected by the added turbulence of the spiral flow and at the same time the effective length of the combustion chamber is increased by the greater length of the spiral combustion flame the combustion zone is thereby able to accommodate. This arrangement thus' permits, in effect, a shorter length of combustion chamber for a given degree of mixing and heat generating capacity than would otherwise be possible. The resultant heated products of combustion and excem air under the combusproximately 6,500 revolutions per minute. Under normal operating conditions the air discharged from the compressorC and entering the combustion zone is maintained at a pressure of aption zonepressure of about '75 pounds per square inch absolute are discharged from the combustion chamber through the channel Hand outward through the nozzle vanes 82 and thence pass substantially radially outward through the expansion zone of the gas turbine G whereindependent counter-rotative impulsion is given to the adjacent rotors of the plurality of concentric turbo-compressor rotors 32 to 31 as before mentioned. The partially expandedgasesexhausted from the gas turbine at I9 after passing through the annular channel l3 formed between the inner proximately pounds per square inch ab olut 76 and outer housing members Hz and H4 are finally discharged rearwardly from the nozzle N in the through the interconnecting space, turbine nozzle vanes 62, and into the annular space 68 adjacent the inner turbine ring 31. The air thus entering the space 68 under pressure passes outwardly substantially radially through the space 69 formed between the rear faces of the rear end rings of the turbine rotors and the annular baiile 13 for the purpose of cooling these rings. A portion of the air thus flowing outward through the space 69 enters the vent holes 59 and 60 as shown in Figure leading into the hollows of each of the turbine blades. A portion of the air thus flowing outward through the space 69 enters each of the innermost vent holes 60 leading through the rings into the hollows of the turbine impeller blades, flows through the channels 56 and 51 formed on either side of the baflies 55 and outward therefrom through the outermost vent holes 59. This flow is induced mainly by the centrifugal force of the air contained within thehollows of the turbine impeller blades under the high rotative speeds which they normally maintain. The cooling air thus having extracted heat from the turbine impeller blades and rings passes around the outer edge of the baflle to return inwardly through the annular channel '10 and from there to discharge forwardly through the slots 1| into the channel 12 leading from the combustion, zone Z to the turbine inlet.

Each of the plurality of individual rows of compressor impeller blades being attached to and carried by independently journaled rotors are thus independently free to select rotational velocities which are determined by the local conditions therein, thus any variations in airplane speeds, altitude or atmospheric temperature will automatically result in optimum adjustment of the ratio of the relative rotative speeds of each adjacent concentric stage or cascade rotor to produce the highest internal turbo-compressor efiiciency of the power plant in the presence of such -external variable conditions. While the tangential speeds of the impeller blades of the several concentric rotors ordinarily maintain themselves at approximately 800 feet per second under the conditions hereindescribed, it may be advantageous to allow the innermost rotors to operate at a tangential speed slightly higher than that of the outermost rotor, but generally not to exceed about greater speed, because in the compressor and turbine flow passages the air and combustion gas temperatures are greatest at the innermost rotors. Under such conditions slightly higher speeds of the inner impeller blades in the higher temperature region may be maintained, and yet remain safely within the allowable Mach number or critical speed which produces compressibility efiects, which eflects are a function of fluid temperature. The pressure ratio through each compressor stage can thus be maintained at a maximum value. This operatingcondition serves to permit almost identical angles of incidence of the blading to exist in all of the rotors'when the solidity of the blading and the lengths of the blades in all of the rotors is substantially identical.

The form of compressor and turbine blading aiiorded by the freely floating squirrel-cage rotors permits the solidity of blading in each rotor, that Y 12 is the ratio of chordal length of the blades to the circumferential spacing between adjacent blades to be greater without deleterious airfoil cascade interference eflects. As a result the ratio of compression of each stage of the compressor may be much higher than in conventional machines without attendant loss of efliciency, which aflords a more efllcient operating cycle with fewer stages, and likewise in the turbine more favorable shapes can be obtained in the combustion gas flow channels defined by the impeller blades. The solidity as defined above may be maintained substantially constant from stage to stage which cannot be accomplished in a conventional machine without entailing other disadvantages well known to those skilled in the art.

It is to be understood that the foregoing is not to be limited but may includeany and all forms of methods and apparatus which are included within the scope of the claims.

I claim:

1. In a turbine apparatus, a plurality of concentric rows of longitudinally positioned peripherally spaced impeller blades of substantially equal length, rings at the opposite ends of said rows, said rings defining a radial flow channel of substantially uniform axial width throughout, said blades of adjacent rows being formed for counter-rotation with respect to one another under the reactive force of radially outward fluid flow through said channel, a separate journal means for only one end ring of each row of blades cantilevering the respective row of blades and its end rings for rotation independently of each other row of blades, and means to pass fluid under pressure radially outward through said channel to drive said blading rows.

2. In a turbine apparatus, a plurality of concentric rows of longitudinally positioned peripherally spaced impeller blades of substantially equal length, rings at the opposite ends of said rows, said rings defining a radial flow channel of substantially uniform axial width throughout, said blades of adjacent rows being formed for counter-rotation with respect to one another under the reactive force of radially outward fluid flow through said channel, journal means supporting one end ring of each row of blades against both radial and axial thrusts, the other end rings of the rows of blades being without journals so that each row of blades is cantilevered from its said one ring, and means to pass fluid under pressure radially outward through said channel to drive said blading rows.

3. In a turbine apparatus, a plurality of concentric rows of longitudinally positioned peripherally spaced impeller blades of substantially equal length, rings at t e opposite ends of said rows, said rings of the several rows defining a radial flow channel of substantially uniform axial width throughout, the blades of adjacent rows being formed for counter-rotation with respect to one another for reactive force upon fluid flowing radially inward through said flow channel, a separate bearing for one ring of each said rows of blades constituting the sole means for supporting the respective row of blades for rotation, each bearing being substantially axially aligned with its respective row of blades, and-means to pass fluid substantially radiallyinward through said flow channel.

4. In a turbine apparatus, a plurality of independently journaled concentric rows of longitudinally positioned peripherallyspaced impeller blades of substantially equal length, an annular shaped flow channel of uniform width axially throughout the radial dimension thereof housing said rows of blades, the said blades of adjacent rows being formed for counter-rotation with respect to one another for reactive force upon fluid flowing radially inward therethrough, means to pass fluid under pressure substantially radially inward through said concentric rows of blades in said annular flowchannel, and a separate bearing means at only one end of each row of blades for cantilevering the respective row of blades for independent rotation.

5. In a turbine apparatus, a plurality of concentric radially intercommunicating rotors each comprising an end ring and a plurality of peripherally spaced longitudinally positioned blades c'arriedby the ring, the blades of adjacent rotors being formed for counter-rotation with respect to one another for reaction with fluid flowing radially through the rotors, a stationary diaphragm, and a separate bearing means supporting the end ring of each rotor on said diaphragm so that the respective rotor has cantilevered support for independent rotation.

6. In a turbo-compressor apparatus, a plurality ofconcentrically positioned radially intercom-- municating squirrel-cage rotors in surrounding relation to one another, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller blades terminating in a common annular ring and bearing means associatedwith each of said rotor rings for rotationally supporting said rotors for independent rotation with respect to one another, each bearing means constituting the sole means for supporting tionally supporting each rotor for independent.

rotation.

8. In a turbo-compressor apparatus, a plurality of concentrically positioned radially intercommunicating squirrel-cage rotors, each of said rotors having a plurality of peripherally. spaced longitudinally positioned impeller blades terminating in a common annular ring, said im eller blades in adjacent concentric rows having pposubstantially radially through said rotors, and a bearing means associated with each of said rotor rings for independently supporting the same so that each of said rotors is cantilevered for independent rotation. I

9. In a turbine apparatus, a plurality of independently journaled concentric rows of longitudinally positioned peripherally spaced airfoil shaped impeller blades of substantially equal length, said impeller blades, in .each row terminating in common end ringsto form unitary cage like rotors and an annular flow channel of substantially uniform width therethrough defined by said annular end rings, said blades of adjacentrows being formed for counter-rotation with respect to one another under the reactive force of outward radial flow of fluid therethrough, means to pass fluid under pressure radially outward through said concentric rows of blades in said annular flow-channel, and stator blades at an end of said flow channel, at least one stator blade having an axially extending duct for equalizing the fluid pressure on opposite externalend surfaces of said end rings whereby axial thrust of said rotor is substantially eliminated.

10. In a turbo-compressor apparatus, a plurality Of concentrically positioned radially interco'mmunic'ating rotors in surrounding relation to one another, each of said rotors having a plu- ,'rality of peripherally spaced longitudinally positioned impeller elements and each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement, said blades terminating in common annular end rings to form a unitary squirrel-' cage rotor, said blades in adjacent concentric rows having opposite angles of incidence for counter-rotation with respect to one another, means to conduct fluid substantailly radially through said rotors, and a bearing means associated with a ring of each rotor cantilevering the rotor for independent rotation. Y

11. In a turbo-compressor apparatus, a plurality of rotors arranged in surrounding relation to one another to have radial intercommunication, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements joined in a common annular end ring site angles of incidence for counter-rotationwith respect to one another,means to conduct fluid to form a unitary squirrel-cage rotor for radial flow of fluid therethrough, and each of said elements having a turbine blade portion] and a compressor blade portion in unitary end to end arrangement, said blades in adjacent concentric rows having opposite angles of incidence for counter-rotation with respect to' one another, means to conduct fluid substantially radially through said rotors, and a bearing means associated with the ring of each rotor for cantilevering the rotor for independent rotation with respect to the other rotors.

12. In a turbo-compressor apparatus, a plurality of rotors arranged-in surrounding relation to one another to have radial intercommunication,

each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements terminating in a common annular end ring to form a unitary squirrel-cage rotor for radial flow of fluid therethrough, and each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement, said blades in adjacent concentric rotors, and a bearing means at the rin of each rotor for cantilevering the rotor forindependent rotation with respect to the other rotors.

. f 13. In a gas reaction propulsive unit, an outer rearward side of said diaphragm, said turbo-compressor having a compressor section and a turbine section, means for rotatably mounting the turbo-compressor on said diaphragm, a combustion chamber in the rearward portion of said inner housing interconnecting said compressor 14. In a gas reaction propulsive unit, a housing having a forward axial opening for the entrance of airv and a rearwardly directed outlet nozzle for the discharging of gases in the form of a propulsive jet, a turbo-compressor in said housing, said turbo-compressor comprising a plurality of concentrically positioned radially intercommunicating rotors, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements, and each of said elements having a turbine blade portion and a compressor blade portion with all of the coznpressor blades of the same length and all of the turbine blades of the same length, forming thereby separate annular compressor and turbine flow channels of substantially uniform width throughout their radial dimensions, a combustion chamber substantially concentrically related to the axis of said rotors, an'outer peripheral inlet to the compressor flow channel communicating with said forward inlet opening for flow of air radially inward through the concentric compressor blades in said compressor flow channel, an inwardly directed discharge from said compressor flow channel leading to said combustion chamber, an outwardly directed radial inlet to said turbine flow channel from said combustion chamber for expansive flow of gases radially outward through said turbine flow channel, and an annular passage leading from the periphery of said gas turbine flow channel and communicating with said rearwardly directed outlet nozzle.

15. In a gas reaction propulsive unit, a housing having a forward axial opening for the entrance of rammed air and a rearwardly directed outlet nozzle for the discharge of gases in the form of a propulsive jet, a turbo-compressor in an intermediate portion of said housing, said turbocompressor comprising a plurality of concentrically positioned independently journaled radially intercommunicating rotors, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements, each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement and terminating in common annular rings which define separate annular shaped compressor and turbine flow channels of a uniform width throughout their radial dimensions, a substantially centrally located combustion chamber, an outer peripheral inlet to the compressor flow channel communicating with said forward inlet opening for flow of air radially inward through the said compressor flow channel, an inwardly directed discharge from said compressor flow channel leading to said central combustion chamber, an outwardly directed radial inlet to said turbine flow channel from said combustion chamber for expansive flow of gas radially outward through said turbine flow-channel, and an annular passage leading from the periphery of said gas turbine flowchannel and communinozzle.

16. In a gas reaction propulsive unit, a housing having a forward axial opening for the entrance of rammed air and a rearwardly directed outlet nozzle for the discharge of gases in the form of a propulsive Jet, a turbo-compressor in an inter mediate portion of said housing, said turbocompressor comprising a plurality of concentrically positioned independently Journaled radially intercommunicating rotors, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements, each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement and terminating in common annular ringsv which define separate annular shaped compressor and turbine flow channels of uniform width throughout their radial dimen-, sions, a substantially centrally located combustion chamber, an outer peripheral inlet to the compressor flow channel communicating with said forward inlet opening for flow of air radially inward through the said compressor flow channel, an inwardly directed discharge from said compressor flow channel, a centrally located conduit leading axially to said annular combustion chamber surrounding said conduit, an outwardly directed radial inlet to said turbine flow channel from said combustion chamber for expans ve flow of as radially outward through said turbine flow channel, and an annular passage leading from the periphery of said gas turbine flow channel and communicating with said rearwardly directed outlet nozzle.

17. In a gas reaction propulsive unit, an outer housing having a forward axial opening for the entrance of rammed air and a rearwardly directed outlet nozzle for the discharge of gases in the form of a propulsive jet, a turbo-compressor in an intermediate portion of said inner housing, said turbo-compressor comprising a plurality of concentrically positioned independently journaled radially intercommunicating rotors, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller 'elements, each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement and terminating in common annular rings which define separate annular shaped compressor and turbine flow channels of uniform width throughout their radial dimensions, a combustion chamber in said inner housing in substantially concentric relation to the axis of said rotors, an outer peripheral inlet to the compressor flow channel communieating with said forward inlet opening for flow of air radially inward through the said compressor flow channel, an inwardly directed discharge from said compressor flow channel, a

expansive flow of gas radially outward through said turbine flow channel, and an annular passage formed between said inner and outer housings extending from theperipheryof said gas turbine flow channel and communicating with said rearwardly directed outlet nozzle.

18. In a gas reaction propulsive unit, an outer housing having a forward axial opening for the entrance of rammed air "and a rearwardly di- 'cating with "said rearwardly directed outlet rected outlet nozzle for the discharge of gases in the form of a propulsive jet, an inner housing within said outer housing, a turbo-compressor in an intermediate portion of said inner housing. said turbo-compressor comprising a plurality of concentrically positioned independently journaled radially intercommunicating rotors, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements, each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement and terminating in common annular rings which define separate annular shaped compressor and turbine flow channels of uniform width throughout their radial dimensions, a combustion chamber in said inner housing in substantially concentric relation to the axis of said rotors, an outer peripheral inlet to the compressor flow channel communicating through an annular passage formed between said inner ,and outer housings with said forward inlet opening for flow of air radially inward through the said compressor flow channel, an inwardly directed discharge from said compressor flow channel, a centrally located conduit leading axially rearward to said combustion chamber, an outwardly directed radial inlet to said turbine flow channel, an annular conduit leading forward from said combustion chamber to said turbine inlet for expansive flow of gas radially outward through said turbine flow channel, and an annular passage formed between said inner and outer housings extending from the periphery of said gas turbine flow channel and communicating with said rearwardly directed outlet nozzle.

19. In a gas reaction propulsive unit according to claim 16 and an accessory compartment in the said inner housing ahead of said turbo-compressor,land drive means extending from said turbocompressor into said accessory compartment.

20. In a turbo-compressor apparatus according to claim 12 in which each bearing means comprises "a track in the rotor ring and slipper bearings received in'the track.

21. Ina gas reaction propulsive unit, a housin having a forward axial opening for the entrance of rammed airand a rearwardly directed outlet nozzle for the discharge of gases in the form of a propulsive jet, a turbo-compressor in an intermediate portion of said housing, said turbocompressor comprising a plurality of concentrically positiond independently journaled radially intercommunicating rotors, each of said rotors having a plurality of peripherally spaced longitudinally positioned impeller elements, each of said elements having a turbine blade portion and a compressor blade portion in unitary end to end arrangement and terminating in common annular rings which define separate annular shaped compressor and turbine flow channels of uniform width throughout their radial dimeninward through the said compressor flow channel,

an inwardly directed discharge from said compressor flow channel leading to said central combustion chamber, an outwardly directed radial inlet to said turbine flow channel from said combustion chamber for expansive flow of gas radially outward through said turbine flow channel, a passage leading from the periphery of said gas turbine flow channel and communicating with said rearwardly directed nozzle, and means for facilitating initiation of operation of the propulsive unit includin a starting motor, and a drive between the motor and one of said rotors whereby the motor is operable to drive said rotor to cause flow through said compressor and turgo bine flow channels.

NATHAN, C. PRICE.

REFERENCES CITED The following references are of record in the 28 file of this patent:

UNITED STATES PATENTS Number Name Date 643,938 Brady Feb. 20, 1900 80 693,946 Boyce Feb. 25, 1902 715,441 Vandegrift Dec. 9, 1902 862,017 Riggs July 30, 1907 879,059 Ludewig Feb. 11, 1908 911,662 Ljungstrom et al. Feb. 9, 1909 85 953,241 Thomson Mar. 29, 1910 1,055,308 Benjamins Mar. 11, 1913 1,076,865 Bonom Oct. 28, 1913 1,091,581 Ljungstrom Mar. 31, 1914 1,197,755 Moller Sept. 12, 1916 40 1,213,889 Lawaczeck Jan. 30, 1917 1,257,167 Wiberg Feb. 19, 1918 1,331,313 Bonom Feb. 17, 1920 1,363,315 Dron Dec. 28, 1920 1,433,950 Kenney Oct. 31, 1922 1,462,592 Bentley July 24, 1923 2,228,425 Venderbush Jan. 14, 1941 2,268,929 Dupont Jan. 6, 1942 2,280,835 Lysholm Apr. 28, 1942 2,318,990 Doran May 11, 1943 2,320,391 Wakefield June 1, 1943 2,357,778 Beaven Sept. 5, 1944 1 2,391,770 Grifilth "Dec. 25, 1945 FOREIGN PATENTS Number Country Date 346,865 France Dec. 13, 1904 411,473 France Jan. 11, 1910 OTHER REFERENCES co Ser. No. 367,666, Anxionnaz et al. (A. P. 0.),

sions, a substantially centrally located combustion chamber, an outer peripheral inlet to the pub. May 25, .1943. 

